Apparatus and method of monitoring operating parameters of a gas turbine

ABSTRACT

A component for use in a combustion turbine ( 10 ) is provided that includes a substrate ( 212 ) and a microelectromechanical system (MEMS) device ( 50, 250 ) affixed to the substrate ( 212 ). At least one connector ( 52 ) may be deposited in electrical communication with the MEMS device ( 50, 250 ) for routing a data signal from the MEMS device ( 50, 250 ) to a termination location ( 59 ). A barrier coating ( 216 ) may be deposited on the substrate ( 212 ) wherein the MEMS device ( 50, 250 ) is affixed beneath a surface of the barrier coating ( 216 ). A plurality of trenches ( 142 ) may be formed in the barrier coating ( 216 ) at respective different depths below the surface of the barrier coating ( 216 ) and a MEMS device ( 50, 250 ) deposited within each of the plurality of trenches ( 142 ). A monitoring system ( 30 ) is provided that may include a processing module ( 34 ) programmed for receiving data from the MEMS device ( 50, 250 ).

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation-in-part of pending U.S. patentapplication Ser. No. 11/122,566 filed May 5, 2005, which claims thebenefit of Provisional Patent Application No. 60/581,662 filed on Jun.21, 2004, which is also a continuation-in-part of U.S. patentapplication Ser. No. 11/018,816 filed Dec. 20, 2004, now U.S. Pat. No.7,270,890 which is a continuation-in-part of U.S. patent applicationSer. No. 10/252,236 filed Sep. 23, 2002, now U.S. Pat. No. 6,838,157 allof which are incorporated herein by reference.

FIELD OF THE INVENTION

The present invention relates generally to monitoring parameters ofoperating environments and particularly to an apparatus and method ofdetermining wear behavior of an abradable coating system deposited oncomponents within an operating environment such as a gas turbine engine.

BACKGROUND OF THE INVENTION

Gas combustion turbines are used for a variety of applications such asdriving an electric generator in a power generating plant or propellinga ship or an aircraft. Firing temperatures in modern gas turbine enginescontinue to increase in response to the demand for higher efficiencyengines. Superalloy materials have been developed to withstand thecorrosive high temperature environment that exists within a gas turbineengine. However, even superalloy materials are not able to withstandextended exposure to the hot combustion gas of a current generation gasturbine engine without some form of cooling and/or thermal insulation.

Thermal barrier coatings are widely used for protecting various hot gaspath components of a gas turbine engine. The reliability of suchcoatings is critical to the overall reliability of the machine. Thedesign limits of such coatings are primarily determined by laboratorydata. However, validation of thermal barrier coating behavior whensubjected to the stresses and temperatures of the actual gas turbineenvironment is essential for a better understanding of the coatinglimitations. Such real world operating environment data is verydifficult to obtain, particularly for components that move during theoperation of the engine, such as the rotating blades of the turbine.

Despite the extreme sophistication of modem turbine engines, such as gasturbines for generating electrical power or aircraft engines forcommercial and military use, designers and operators have very littleinformation regarding the internal status of the turbine enginecomponents during operation. This is due to the harsh operatingconditions, which have prevented the use of traditional sensors forcollecting reliable information of critical engine components.

Many current turbines are equipped with sensors capable of limitedfunctions such as exhaust gas-path temperature measurements, flamedetection and basic turbine operating conditions. Based on thisinformation, turbine operators such as utility companies operate enginesin a passive mode, in which maintenance is scheduled based on priorhistories of similar engines. Engine rebuilds and routine maintenanceare performed in the absence of a prior knowledge of the remaining oralready utilized life of individual components. The lack of specificcomponent information makes early failure detection very difficult,often with the consequence of catastrophic engine failure due to abruptpart failure. This results in inefficient utilization, unnecessarydowntime and an enormous increase in operating cost.

Currently, the gas turbine industry approach is to depend on themeasurement of gas path temperature, which is related back to specificcomponent problems based on experience and history regarding a class ofengines. This approach is highly subjective and only allows fordetermining already severe situations with an engine. It does notprovide indications of impending damage or insight into the progressionof events leading up to and causing engine damage due to componentdegradation or failure.

The instrumentation of a component such as a blade or vane within asteam turbine typically includes placing wire leads on the balancewheel, which continue on to the blade airfoil. The wire leads aretypically held together by an epoxy. These wires are routed from withinthe component to the turbine casing. The pressure boundary of acomponent may be breached to introduce a sensor such as a thermocoupleand a braze is back filled to hold the thermocouple in place. Eachthermocouple sensor has wire leads coming out of the component that areconnected back to a diagnostic unit. Instrumenting a plurality ofcomponents of a turbine in this manner results in an extensive networkof wires just for monitoring the single operating condition oftemperature. Instrumenting components using this technique is expensive,which is a barrier to instrumenting a large number of components withina single turbine. Further, the wire leads and data transfer isfrequently poor, which can result in costly repairs and flawed dataanalysis.

Using thermocouples for temperature measurements in the gas path of aturbine may be disadvantageous because it only provides feedback to anoperator that a temperature change has occurred in the gas path. It doesnot provide any indication as to why the temperature change hasoccurred. For diagnosing problems with blades or vanes based on ameasured temperature change, there has to be an historical correlationbetween the measured temperature differential and the specific problem,such as a hole in a vane. This correlation is difficult and timeconsuming to derive to within a reasonable degree of certainty and needsto be done on an engine-by-engine basis taking into account turbineoperation conditions. When a temperature differential is measured, it isdifficult, if not impossible, to predict what the problem is or where itis located. Consequently, the turbine must typically be shut down andinspected to determine the scope of repair, replacement or othermaintenance to be performed.

In any application, combustion turbines are routinely subject to variousmaintenance procedures as part of their normal operation. Diagnosticmonitoring systems for gas turbines commonly include performancemonitoring equipment that collects relevant trend and fault data usedfor diagnostic trending. In diagnostic trend analysis, certain processdata (such as exhaust gas temperature, fuel flow, rotor speed and thelike) that are indicative of overall gas turbine performance and/orcondition are compared to a parametric baseline for the gas turbine. Anydivergence of the raw trend data from the parametric baseline may beindicative of a present or future condition that requires maintenance.Such diagnostic monitoring systems can only predict or estimate specificcomponent conditions and do not collect data from or provide anyanalysis with respect to the actual condition of a specific componentitself.

In this respect, conventional methods of predicting component failurefor gas turbines and of scheduling maintenance have not been entirelyaccurate or optimized. The traditional “duty cycle” used for predictivemaintenance does not reflect real operational conditions, especiallyoff-design operations. The actual life of specific components of a gasturbine depends strongly on the actual usage of that gas turbine and thespecific components within the turbine.

For example, elevated temperatures and stresses within the turbine, andaggressive environmental conditions may cause excessive wear oncomponents in the turbine beyond that predicted with the standard designduty cycle. Off-design operating conditions, which are often experiencedby industrial gas turbines, are not reflected by the standard dutycycles. The actual life of components in the gas turbine may besubstantially less than that predicted by the design duty cycle.Alternatively, if more favorable conditions are experienced by an actualgas turbine than are reflected in the design duty cycle, the actualcomponent life may last substantially longer than that predicted bymaintenance schedules based on the design duty cycle. In either event,the standard design duty cycle model for predicting preventivemaintenance does not reliably indicate the actual wear and tearexperienced by gas turbine components.

Known techniques for predicting maintenance and component replacementrely on skilled technicians to acquire or interpret data regarding theoperation of a combustion turbine. Such techniques are subject tovarying interpretations of that data by technicians. Technicians maymanually evaluate the operational logs and/or data collected from gasturbines. Technicians, for example, may evaluate start and stop timesand power settings to determine how many duty cycles had beenexperienced by the gas turbine, their frequency, period and otherfactors. In addition, if the data log of a gas turbine indicated thatextraordinary conditions existed, such as excessive temperatures orstresses, the technicians may apply “maintenance factors” to quantifythe severity of these off-design operational conditions.

None of these techniques provide accurate information with respect tothe actual condition of a specific component or component coating, whichmay lead to unnecessary repair, replacement or maintenance beingperformed causing a significant increase in operating costs.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross sectional view of an exemplary combustion turbine withwhich embodiments of the invention may be used and an exemplarymonitoring and control system for collecting and analyzing componentdata from the combustion.

FIG. 2 a perspective view of an exemplary combustion turbine vaneequipped with an exemplary embodiment of the present invention.

FIG. 3 is a schematic view of a vane of FIG. 2.

FIG. 4 is a schematic cross section of the compressor of FIG. 1.

FIG. 5 is a perspective partial view of an exemplary embodiment of asmart component combustion in accordance with aspects of the invention.

FIG. 6A is a schematic view of an exemplary embodiment of the componentof FIG. 5.

FIG. 6B is a schematic view of an exemplary embodiment of the componentof FIG. 5.

FIG. 6C is a schematic view of an exemplary embodiment of the componentof FIG. 5.

FIG. 7 is an exemplary embodiment of a heat flux sensor.

FIGS. 8 and 9 illustrate an exemplary embodiment of a strain gauge and acrack propagating to different lengths.

FIG. 10 is a partial perspective view of a component having a sensorembedded within a layer of thermal barrier coating material disposedover a substrate material.

FIG. 11 is a partial cross-sectional view of a component having aplurality of sensors embedded at varying depths below a surface of thecomponent.

FIG. 12 is a process diagram illustrating steps in a method ofmanufacturing the component of FIG. 11.

FIG. 13 is a partial cross-sectional view of a component having aplurality of sensors embedded at varying depths below a surface of thecomponent.

FIG. 14 is schematic plan view of an exemplary microelectromechanicalsystem (MEMS) device.

FIG. 15 is a perspective view of exemplary MEMS device embedded in anabradable coating system.

DETAILED DESCRIPTION OF THE INVENTION

Embodiments of the present invention may use microelectromechanicalsystems (MEMS) devices as sensors embedded within various types ofcoatings recognized by those skilled in the art. For example, barriercoating may be used herein generally to refer to a range of coatingscommonly used in combustion turbine engines such as abradable coatingsystems, thermal barrier coatings, CMC coatings, wear coatings,protective overlay coatings, insulating coatings and restorationcoatings as well as others. Reference to specific types of coatingsherein is by way of example only.

MEMS devices may be embedded in barrier coatings and/or affixed on orwithin a surface of components to enable monitoring and diagnostics of asystem such as an exemplary combustion turbine 10 of FIG. 1. Using MEMSdevices is advantageous because they may be placed directly at locationsof interest due to their small size and robust electrical connections.Locating MEMS devices directly at locations of interest provides anincreased accuracy in measurements relative to remote sensors that arelocated away from the locations of interest, in which case measurementsmust be extrapolated to predict events at the location of interest. MEMSdevices may be coupled with antenna located on their respective siliconchips for wireless transmission of data indicative of the desiredproperties being measured or monitored.

FIG. 1 illustrates an exemplary combustion turbine 10 such as a gasturbine used for generating electricity as will be recognized by thoseskilled in the art. Embodiments of the invention may be used withcombustion turbine 10 or in numerous other operating environments andfor various purposes as will be recognized by those skilled in the art.For example, embodiments may be used in aircraft engines, monitoringtemperature and heat flux in boilers, heat exchangers and exhauststacks; determining insulation performance and degradation; determiningpipe fouling; and evaluating vibrating component health. Embodiments maybe used in the automotive industry for monitoring combustion chamberconditions, rotating components such as crankshaft, cams, transmissionsand differentials, and determining suspension and frame integrity forheavy-duty vehicles. Embodiments may also be used in measuring strainand heat flux in tanks, portable and other equipment operating indessert, wet, and/or high temperature configurations.

Returning to FIG. 1, combustion turbine 10 includes a compressor 12, atleast one combustor 14 (broken away) and a turbine 16. Compressor 12,combustor 14 and turbine 16 are sometimes referred to collectively as agas turbine engine. Turbine 16 includes a plurality of rotating blades18, secured to a rotatable central shaft 20. A plurality of stationaryvanes 22 are positioned between blades 18, with vanes 22 beingdimensioned and configured to guide air over blades 18. Blades 18 andvanes 22 will typically be made from nickel-cobalt, and may be coatedwith a thermal barrier coating 26, such as yttria-stabilized zirconia.Similarly, compressor 12 includes a plurality of rotating blades 19positioned between respective vanes 23.

In use, air is drawn in through compressor 12, where it is compressedand driven towards combustor 14. Combustor 14 mixes the air with fueland ignites it thereby forming a working gas. This working gas willtypically be above 1300° C. This gas expands through turbine 16, beingguided across blades 18 by vanes 22. As the gas passes through turbine16, it rotates blades 18 and shaft 20, thereby transmitting usablemechanical work through shaft 20. Combustion turbine 10 may also includea cooling system (not shown), dimensioned and configured to supply acoolant, for example steam or compressed air, to blades 18 and vanes 22.

The environment wherein blades 18 and vanes 22 operate is subject tohigh operating temperatures and is particularly harsh, which may resultin serious deterioration of blades 18 and vanes 22. This is especiallylikely if the thermal barrier coating 26 should spall or otherwisedeteriorate. Embodiments of the invention are advantageous because theyallow components to be configured for transmitting data indicative of acomponent's condition during operation of combustion turbine 10. Blades18, 19, vanes 22, 23, and coatings 26, for example, may be configuredfor transmitting component specific data that may be directly monitoredto determine the respective condition of each component during operationand to develop predictive maintenance schedules.

FIG. 1 also illustrates a schematic of an exemplary monitoring andcontrol system 30 that may be used in accordance with various aspects ofthe present invention. System 30 may include an antenna 32, a receiver33, a processor or CPU 34, a database 36 and a display 38. Processor 34,database 36 and display 38 may be conventional components and antenna 32and receiver 33 may have performance specifications that are a functionof various embodiments of the invention. For example, antenna 32 andreceiver 33 may be selected for receiving wireless telemetry datatransmitted from a plurality of transmitters deployed in variouslocations throughout combustion turbine 10 as more fully describedbelow.

Embodiments of the present invention allow for a plurality of MEMSsensors to be embedded within the respective coatings of a plurality ofcomponents within combustion turbine 10. Alternate embodiments allow forthe sensors to be surface mounted or deposited to components, especiallythose contained in areas where components do not require a barriercoating such as the compressor. Exemplary embodiments of sensors may beused to provide data to system 30 with respect to physicalcharacteristics of a component and/or properties of a component'scoating as well as other component or coating specific information.

For example, exemplary sensors may be used to detect wear between twocomponents, measure heat flux across a component's coating, detectspalling of a coating, measure strain across an area of a component ordetermine crack formation within a component or coating. MEMS sensorsmay be configured as proximity probes, accelerometers, load cells,pressure transducers, strain gauges, temperature probes, heat fluxsensors, vibration sensors and gas sensors. Those skilled in the artwill recognize other properties and/or characteristics of a component,component coatings and operating parameters of combustion turbine 10that may be monitored, measured and/or detected in accordance withaspects of the invention.

It will be appreciated that aspects of the invention allow for variousMEMS sensor configurations to be embedded within a barrier coating suchas a barrier coating 26 of blades 18 or vanes 22 of turbine 16. U.S.Pat. No. 6,838,157, which is specifically incorporated herein byreference, describes various embodiments of methods for instrumentinggas turbine components, such as blades 18 and vanes 22 that may beutilized for depositing MEMS sensors in accordance with aspects of thepresent invention. This patent discloses various methods of formingtrenches in a barrier coating, forming a sensor in the coating anddepositing a backfill material in the trench over the coating.Embodiments of those methods and components may be used to form smartcomponents incorporating MEMS sensors as disclosed herein.

U.S. Pat. No. 6,576,861, which is specifically incorporated herein byreference, discloses a method and apparatus that may be used to depositembodiments of sensors and sensor connectors with transmitters inaccordance with aspects of the present invention. In this respect,methods and apparatus disclosed therein may be used for the patterningof fine sensor and/or connector features of between about 100 micronsand 500 microns without the need of using masks. Multilayer electricalcircuits and sensors may be formed by depositing features usingconductive materials, resistive materials, dielectric materials,insulative materials and other application specific materials. It willbe appreciated that other methods may be used to deposit multilayerelectrical circuits and sensors in accordance with aspects of theinvention. For example, thermal spraying, vapor deposition, lasersintering and curing deposits of material sprayed at lower temperaturesmay be used as well as other suitable techniques recognized by thoseskilled in the art.

Embodiments of the invention allow for a plurality of sensors 50, whichmay be MEMS devices to be deployed in numerous places within combustionturbine 10 for monitoring component-specific or coating-specificconditions as well as collecting other data with respect to theoperation or performance of combustion turbine 10. For example, FIG. 1illustrates that one or more sensors 50 may be embedded withinrespective barrier coatings 26 of one or more blades 18 of turbine 16.It will be appreciated that sensors 50 may be embedded within barriercoatings of other components with turbine 16 for whichcomponent-specific and/or coating-specific data is to be acquired.

FIG. 2 illustrates a pair of vanes 23 removed from compressor 12 withone vane having a sensor 50 mounted or connected with vane 23 fordetecting a condition of vane 23. A connector 52 may be provided for asa means for routing a data signal from sensor 50 to a transmitter 54configured for wirelessly transmitting the data signal to a transceiver56. Connector 52 may be one or a plurality of electrical leads forconducting a signal from sensor 50 to a surface mounted transmitter 54.Alternate embodiments allow for various types of connectors 52 to beused as a means for routing a data signal from sensor 50 to transmitter54, depending on the specific application. For example, one or aplurality of fiber optic connectors may be used for routing a signalusing single or varying wavelengths of light.

Embodiments allow for transmitters 54 to be multi-channel and havevarious specifications depending on their location within a casing ofcombustion turbine 10. Transmitters 54 may be configured to functionwithin the compressor 12 casing subject to operating temperatures ofbetween about 80° C. to 120° C. They may also be configured to functionwithin the turbine 12 casing subject to operating temperatures ofbetween about 300° C. to 350° C. of higher, and be resistant tooxidative exposure.

FIG. 3 illustrates a schematic plan view of compressor vane 23 havingsensor 50 connected therewith and connector 52 connecting sensor 50 withtransmitter 54. A power source 51 may be provided, such as anappropriately sized battery for powering transmitter 54. In alternateembodiments transmitter 54 may be located remotely from vane 23 andpowered from an external power source. Transmitter 54 may receivesignals from sensor 50 via connector 52 that are subsequently wirelesslytransmitted to transceiver 56. Transceiver 56 may be mounted on hub 58or on a surface external to compressor 12 such as the exemplarylocations shown in FIG. 1. Transceiver 56 may be mounted in variouslocations provided it is within sufficient proximity to transmitter 54to receive a wireless data transmission, such as an RF signal fromtransmitter 54. Transceiver 56 may transmit the RF signal to antenna 32of system 30 where the signal may be processed for monitoring thecondition of compressor vane 23.

With respect to FIGS. 2 and 3, one or more sensors 50 may be connectedwith one or more compressor vanes 23 by fabricating sensor 50 directlyonto a surface of vane 23. Connector 52 may be deposited directly onto asurface of vane 23. In alternate embodiments a trench or recess may beformed within a surface of vane 23 that is sized for receiving adeposited sensor 50 and connector 52. Sensor 50 and connector 52 may bedeposited within the recess and protected by depositing a coating ofsuitable material onto a surface of vane 23 over sensor 50 and connector52. In other alternate embodiments a coating may be deposited onto asurface of vane 23, a trench may be formed within the coating and sensor50 and connector 52 may be deposited within the trench. A protectivecoating may be deposited over sensor 50 and/or connector 52.

Connector 52 may extend from sensor 50 to a termination location, suchas the peripheral edge of vane 23 so that a distal end 53 of connector52 is exposed for connection to transmitter 54. Sensor 50 and connector52 may be positioned on vane 23 to minimize any adverse affect on theaerodynamics of vane 23.

In an embodiment, one or more sensors 50, such as strain gauges orthermocouples, for example, may be deposited on one or more turbine orcompressor blades 18, 19. FIG. 4 illustrates an embodiment with respectto compressor 12. A connector 52 may be deposited to connect each sensor50 to one or more transmitters 54 connected with blade 18, 19. It willbe appreciated that exemplary embodiments allow for a plurality ofsensors 50 to be connected with a single transmitter 54 via respectiveconnectors 52. For example, a sensor 50 may be deposited on each of aplurality of blades 18, 19. A connector 52 may be deposited to route asignal from each sensor 50 to a single transmitter 54.

Transmitter 54 and a rotating antenna 55 may be mounted proximate theroot of blade 18, 19. Connector 52 may be routed from sensor 50 aft tothe root of blade 18, 19 to connect sensor 50 with rotating antenna 55,which may in turn be connected with transmitter 54 via a connector 52 a.A stationary antenna 57 may be installed on a turbine or compressor vane22, 23 aft of the root of respective blade 18, 19. A lead wire 57 a maybe routed from stationary antenna 57 out of compressor 12 or turbine 16to broadcast a signal to system 30. In exemplary embodiments, such asthat shown in FIG. 4, power may be generated through induction duringoperation of compressor 12 as will be appreciated by those skilled inthe art. In this arrangement, transmitter 54 may transmit data tostationary antenna 57 via rotating antenna 55 and power may be suppliedfrom stationary antenna 57 to transmitter 54.

It will be appreciated by those skilled in the art that one or moresensors 50 may be mounted to, such as by a spray deposition, eachcompressor blade 19 within a row of blades 19 mounted on a disk withincompressor 12. A respective connector 52 may connect each sensor 50 to arespective transmitter 54 mounted proximate the root of each blade 19within the row. Rotating antenna 55 may encircle the disk proximate theroot of each blade 19 and be connected with each transmitter 54 via arespective connector 52 a. One or more stationary antennas 57 may beinstalled on a compressor vane 23 aft of the row of compressor blades19, or in another location, such as a compressor hub sufficientlyproximate to rotating antenna 55 for signal broadcasting and receiving.Stationary antenna 57 may also encircle the row of blades 19. Rows ofblades 18 in turbine 16 may be similarly configured.

FIG. 5 illustrates a partial view of a component, such as a vane 22 fromturbine 16 having a barrier coating 26 deposited thereon. Sensor 50 andconnector 52 may be embedded beneath an upper surface of barrier coating26. Connector 52 may have a distal end 53 that is exposed at atermination location, such as proximate a peripheral edge 59 of vane 22for connection with transmitter 54. In an embodiment transmitter 54 maybe surface mounted to vane 22 or embedded within coating 26 proximateperipheral edge 59. Alternate embodiments allow for transmitter 54 to belocated elsewhere such as on a platform (not shown) to which vane 22 isconnected or in a cooling flow channel, for example, as will berecognized by those skilled in the art.

FIG. 6A illustrates a schematic plan view of a blade 18 having anexemplary sensor 50 connected therewith and connector 52 connectingsensor 50 with transmitter 54. Transmitter 54 may be powered throughinduction generated within turbine 16 during operation that will beappreciated by those skilled in the art. FIGS. 6A, 6B and 6C illustrateexemplary embodiments of a turbine blade 18 having transmitter 54 placedin various locations. In FIGS. 6A and 6B transmitter 54 may be mountedto blade 18 and FIG. 5C illustrates that transmitter 54 may be locatedremote from blade 18. For example, transmitter 54 may be locatedremotely from blade 18 such as within a disk (not shown) to which aplurality of blades 18 is attached. In this respect, transmitter 54 maybe maintained in a cooler location outside the hot gas path, which mayincrease the transmitter's useful life. Locating transmitter 54 remotefrom blade 18 allows for using an external power source for poweringtransmitter 54 rather than using a battery or induction.

A power supply may also be attached to sensor 50 to provide additionalfunctionality to the sensor. This additional functionality could includemechanical actuation as a result of feedback to the sensor 50 output.Such an integrated system may be applicable for components, such as ringsegments for real-time gap control.

The exemplary embodiments of compressor vane 23 and turbine blade 18illustrated in FIGS. 3-6A, 6B and 6C configured with self-containedsensors 50 and connectors 52 are advantageous in that they may beprefabricated for installation in combustion turbine 10 by a fieldtechnician. Embodiments allow for a distal end 53 of connectors 52 to beexposed at a termination location. This location may be proximate aperipheral edge of a component or other location. This allows a fieldtechnician to quickly and easily connect connector 52 to a transmitter54 regardless of its location.

Providing components of combustion turbine 10, such as vanes 23 and/orblades 18 with pre-installed sensors 50 and connectors 52 is asignificant advantage over previous techniques for installing suchcomponents in the field, which typically required an extensive array ofwires to be routed within combustion turbine 16. Providing componentswith pre-installed sensors 50 and connectors 52 allows for monitoringthe condition of those specific components during operation ofcombustion turbine 10.

Embodiments of the invention allow for sensor 50 to be configured toperform a wide range of functions. For example, sensor 50 may beconfigured to detect wear of a single component or between twocomponents, measure heat flux across a component's coating, detectspalling of a coating, measure strain across an area of a component ordetermine crack formation within a component or coating. U.S. patentapplication having application Ser. No. 11/018,816 discloses embodimentsof a system that generally involves monitoring the wear of a componentthat may be configured in accordance with embodiments of the presentinvention.

Wear sensors 50 may be configured as embedded electrical circuits in acontact surface of a component, such as a tip of blade 18 and thecircuit may be monitored by monitoring system 30 for indications ofwear. By positioning a circuit at the wear limit, or at prescribeddepths from the component's surface, the condition of the surface may becontinuously monitored and system 30 may provide an operator with anadvanced warning of service requirements.

It will be appreciated that sensor 50 may be configured for weardetection and prefabricated within a component for use within combustionturbine 10 either alone or in combination with a means for transmitting52 in accordance with aspects of the present invention. In this respect,the signals extracted for detection of wear may be conducted viaconnectors 52 to transmitter 54, which may transmit the signals viawireless telemetry to a transceiver 56 and subsequently system 30.

Embodiments of the present invention allow for monitoring and controlsystem 30 to collect and store historical data with respect to acomponent's wear and correlating the component's wear with the operatingconditions of combustion turbine 10 responsible for producing the wear.This may be accomplished by continuously interrogating turbine 16conditions, for example, by the deposition of piezoelectric devicesand/or other sensors 50 configured for providing a continuous datastream indicative of the loading conditions and vibration frequencyexperienced by various components within turbine 16. This data may becorrelated to data indicative of a component's wear and used forpredictive maintenance or other corrective actions.

FIG. 7 illustrates another exemplary embodiment of a sensor 50 that maybe configured as an exemplary heat flux sensor 61 for measuring heatflux across a barrier coating such as a thermal barrier coating (TBC)60, which may be yttrium-stabilized zirconium. Using known techniques,thermal barrier coating 60 may be deposited on a bond coat 62, which maybe deposited on a substrate 64. Substrate 64 may be various componentssuch as a superalloy suitable for use in turbine 16, and in anembodiment may a blade 18. The heat flux may be used to obtain thesurface temperature of substrate 64 without having to expose the surfaceof substrate 64 to the surface temperature experienced by the uppersurface of thermal barrier coating 60.

Thermocouples 66 may comprise a material having a coefficient of thermalexpansion that substantially matches that of the material within whichthey are deposited, such as thermal barrier coating 60. In anembodiment, a plurality of temperature sensors, such as K-typethermocouples 66 may be embedded within a thermal barrier coating 60with thermocouples 66 located vertically over each other as shown inFIG. 6. In an embodiment, thermocouples 66 may include a NiCr/NiAlthermocouple junction. Alternate embodiments allow for thermocouples 66to be fabricated of other materials such as Pt and Pt—Rh for hightemperature applications such as those within turbine 16.

Heat flux sensor 61 may be formed in different geometries to achieve adesired signal-to-noise ratio. Each thermocouple 66 may be approximately25 microns thick but this thickness may vary depending on theapplication. Because the thermal barrier coating 60 may be several timesas thick as thermocouples 66 they will not significantly alter theprofile or performance of thermal barrier coating 60. Embodiments allowfor post deposition laser micromachining to achieve a desired junctiondensity.

As heat flows vertically into or out of thermal barrier coating 60, eachthermocouple 66 will record a different temperature measurement. Bymeasuring the temperature differences and knowing the thickness andthermal conductivity of thermal barrier coating 60, the heat flux can beobtained. Thermocouples 66 may be connected with a means fortransmitting 52 as described herein so that the respective temperaturemeasurements taken by each thermocouple 66 may be wirelessly transmittedto monitoring and control system 30.

FIGS. 8 and 9 illustrate an exemplary embodiment of a sensor 50 that maybe configured as an exemplary sensor 68 configured for detecting and/ormeasuring strain or a crack within a location of interest such assubstrate 70. For example, substrate 70 may be a location of interest ofa surface area of a blade 18, or it may be other locations of interestwithin or at the surface of thermal barrier coating 60 or bond coat 62.It will be appreciated that sensor 68 configured in this manner may beused in numerous places throughout combustion turbine 10. The sensorsdescribed in FIGS. 8 and 9 describe the utilization of the change inresistance to result in a strain output. Other embodiments of straingauges could also include capacitive changes to determine the localstrain values.

In this respect, critical engineering components, such as blades 18, 19and vanes 22, 23 are nearly universally subjected to some form ofmechanical and/or thermo-mechanical cyclic loading. Aspects of theinvention allow for the assessment of component service life by theintermittent or continuous, in-situ measurement of applied strains andcrack detection with respect to that component. This may be accomplishedby the placement of embedded strain gages and crack sensors 68 invarious locations within combustion turbine 10. Sensors 50 configured asa strain gauge 68 may be formed using a NiCr material for use in lowertemperature applications, such as in compressor 12 of combustion turbine10.

Sensors 68 may be used as crack sensors by placing them at locations orpoints where cracks are known or likely to appear. A crack sensor gauge68 may be optimized for size, crack propagation, and crack extentthrough appropriate choice of gauge 68 parameters. Such parameters mayinclude the footprint of gauge 68, spacing of fingers 72, andorientation of fingers 72 with respect to the direction of a predictedcrack propagation. Crack formation in substrate 70 gives rise to alarge, abrupt change in the strain gauge response, and may be detectedby continuously monitoring the sensor 68 output for abrupt signalchanges using known signal processing techniques. Data indicative of thesignal change may be conducted via a means for transmitting 54 to atransceiver 56 and subsequently transmitted to monitoring and controlsystem 30 via wireless telemetry.

In an exemplary embodiment, a strain gauge sensor 68 may be bonded to ordeposited on a surface of a compressor blade 19 and positioned so thatbending stress on blade 19 varies the output signal from sensor 68.Connector 52, which may be wire leads, are routed to a transmitter 54located on a rotating collar internal to compressor 12. Transmitter 54may have an onboard bridge completion and provide a regulated voltage tosensor 68. As the output signal from sensor 68 varies an RF signal fromtransmitter 54 varies proportionally. The RF signal may be transmittedto a transceiver 56, which receives the RF signal and converts it into avoltage signal proportional to the strain detected by sensor 68. The RFsignal may be transmitted to system 30. An exemplary transmitter 54 maypick up changes in strain from about 30 Hz to about 30 KHz.

Embodiments of the invention allow for using crack sensors 68 to monitorcrack growth during operation of combustion turbine 10 and verify designmodels by varying component operating parameters until cracks aredetected with the crack sensors 68. The design models will be calculatedfor the same operating parameters to see if they successfully predictcrack growth and formation, and will be modified accordingly.

Monitoring and control system 30 may collect and store data indicativeof strain and crack measurements from numerous components in criticallocations within combustion turbine 10, such as blades 18, for example.Such data may be analyzed over time to develop a strain history for eachcomponent. A component's strain history may include the magnitude andorientation of strains, and the occurrence of overloads under cyclicloading. An appraisal of fatigue damage may be developed and used forpredictive maintenance.

Embodiments of the present invention allow for deploying a plurality ofsensors 50 throughout combustion turbine 10 by either surface mountingthem to components or embedding them within respective component barriercoatings to collect specific component condition data and transmit thatdata using wireless telemetry to monitoring and control system 30. Thisapproach is advantageous in that it allows for the replacement, repairand maintenance decision-making processes to be based on the conditionof specific components during operation of combustion turbine 10.

In this respect, specific component condition data may be received byantenna 32 and receiver 33 then stored in database 36 by CPU 34.Embodiments allow for specific component condition data to be collectedand presented to an operator in real time via display 38. This allowsfor an operator to make instantaneous decisions regarding the operationof combustion turbine 10 in response to the condition of a specificcomponent or components.

Historical data may be compiled and analyzed with respect to eachcomponent for making repair, replacement or maintenance decisions withrespect to that component. Operating conditions and specific componentsof combustion turbine 12 may be monitored sets of conditions may beisolated that are indicative of a component or components needing to berepaired or replaced, or of corrective action to be taken with respectto operation of the gas turbine. These aspects allow for significantimprovement in predictive maintenance schedules.

FIG. 10 is a partial perspective illustration of a component 110 formedof a substrate material 112 having a barrier coating such as a layer ofthermal barrier coating 114 disposed on one surface 116. The component110 may be part of a gas turbine engine 10 of FIG. 1, for example, orany other machine wherein a base material must be protected from anexternal environment by a layer of a barrier material. In an embodiment,component 110 may be an airfoil member, such as a turbine blade 18disposed in the hot gas flow path of a engine 10 with an oxide ornon-oxide ceramic TBC 14 such as mullite, silicon carbide or azirconium-based ceramic overlying a superalloy substrate material 112.

Component 110 may alternatively be fabricated from a ceramic matrixcomposite (CMC) substrate coated with an environmental barrier coating(EBC) or a thermal barrier coating (TBC). Because the integrity of thecoating 114 is critical to the overall integrity of the component 110,it is useful to obtain operating parameter information that directlyaffects the performance of the coating 114. Such information is obtainedby embedding a sensor, such as a sensor 50 below the exposed surface 118of the TBC 114. The sensor is not visible in FIG. 10 but may be locatedbelow surface 118 in the sensing location indicated generally by numeral120.

The sensor may be one that provides a signal indicative of temperature,strain, crack initiation, chemical changes, vibration, pressure or otherparameters of interest. These sensors themselves could be multi-layeredcontaining a combination of electrodes and the functional body.Conductors 122 may also be located below surface 118 may route thesignal produced by the sensor away from sensing location 120 to atermination location, which may be a connection location indicatedgenerally by numeral 224 where they can conveniently exit the component110. Conductors 122 may function similarly to connectors 52 for routinga signal from a sensor, such as a sensor 50 to a transmitter 54 fortransmission to system 30 via wireless telemetry. The sensor and theconductors 122 may be insulated from the surrounding environment by alayer of insulating material 126.

FIG. 11 is a partial cross-sectional view of another component 130having a substrate material 132 covered by a barrier coating such as alayer of a thermal barrier coating material 134 for use in a very hightemperature environment. As is well known in the art of TBC coatings, abond coat 136 such as an MCrAlY material may be deposited on thesubstrate 132 prior to the application of the TBC material 134 toimprove the adherence of the coating 134 to the substrate 132.

Component 130 may be instrumented by a plurality of sensors, such assensors 50 embedded at a plurality of depths below a surface 138 of theTBC material 134 that is exposed to the external environment. A firstsensor 140 is deposited in a relatively shallow trench 142. Trench 142may be lined with an electrically insulating coating 144 such asaluminum oxide to prevent the grounding of sensor 140 to the TBCmaterial 134. Sensor 140 may take any form known in the art, for examplea thermocouple formed by a bimetallic thermocouple junction or othersensors described herein. The surface location of sensor 140 suggeststhat it may be useful for sensing a parameter related to the externalenvironment, such as temperature or a chemical parameter.

FIG. 12 illustrates the steps of a process 150 that may be used duringthe manufacturing of the component 130 of FIG. 11. In step 152, a layerof thermal barrier coating material 134 may be deposited onto asubstrate 132. After step 152, the component is completed in its normaloperating shape as it may be used without embedded instrumentation. Oneskilled in the art may appreciate, therefore, that the process 150 maybe applied to newly fabricated components or it may be back fit to anexisting component that is in inventory or that has been in service.

In step 154, a trench 142 may be formed in a surface 138 of thecomponent 130. Trench 142 may be formed to any desired shape by anyknown method, such as by laser engraving trench 142 to have a generallyrectangular cross-section with a predetermined width and depth.Variables for such a laser engraving process include spot size, powerlevel, energy density, pulse frequency, and scan speed. These variablestogether affect the trench width, depth, material removal rate and thecost of manufacturing. Trench 142 may have a constant cross-sectionalsize and shape along its entire length, or it may vary in size and/orshape from one region to another. For example, in the component 110 ofFIG. 10, a trench formed in the sensing location 120 may have differentdimensions than the trench extending from the sensing location 120 tothe connecting location 124, since the sensor and the conductors 122 mayhave different geometries. The trench 142 may also be inclined to thesurface, i.e. varying in depth along its length, which in someapplications may provide improved mechanical integrity within thecomponent.

After trench 142 is formed at step 154, an insulating coating 144 may beapplied to the surfaces of the trench 142 at step 56 in order to provideelectrical isolation between sensor 140 and TBC material 134. Insulatingcoating 144 may be deposited by any known method such as chemical vapordeposition (CVD) to a thickness sufficient to achieve a desired level ofelectrical isolation. Once the trench 142 is formed at step 154 andinsulated at step 156, the sensor 140 may be formed by depositing theappropriate material or materials into trench 142 at step 158. Any knownmaterial deposition process providing the desired material propertiesmay be used. Such processes are common in the fields of rapidprototyping, thin and thick film deposition, and thermal spraying, andinclude, for example, chemical vapor deposition, plasma spray,micro-plasma spray, cold spray, electroplating, electrophoreticdeposition, HVOF, sputtering, CCVD, sol-gel and selective laser melting.Processes typically used for the fabrication of multi-layer thick filmcapacitors may also be used, such as the application of pastes and tapesof the desired materials.

After the deposition of material, a heat input may be used to sinter thematerial, thereby increasing the mechanical integrity of the sensor.This can be done either by heating using a flame, plasma, furnaceannealing or localized laser energy application. In the selective lasermelting (SLM) process, powdered material having a predeterminedchemistry may be deposited into the trench and melted with the energy ofa laser beam to form the respective portion of the sensor 140 of FIG. 11or the interconnecting conductors 122 of FIG. 10. For example, to form athermocouple, platinum powder may be deposited into one portion oftrench 142 and solidified by a SLM process. Platinum-rhodium powder maythen be deposited into a second portion of trench 142, either along thetrench length or as a second vertical layer, and solidified by a SLMprocess to contact the platinum material to form the thermocouplejunction.

Note that the geometry of trench 142 may have a direct effect on thegeometry of the sensor 140. Accordingly, it is possible to affect theoperating parameters of sensor 140 or interconnecting conductors 122 bycontrolling the dimensions of the respective trench 142. For example,the resistance of a conducting line formed within a trench will beaffected by the width of the trench. The laser engraving process of step154 may be closely controlled to achieve a desired trench geometry.Certain commercially available processes for depositing a conductor ontoa flat surface by thermal spraying may not produce the fine featuresthat may be necessary for sensors and conductive lines. Such processesmay rely on a subsequent material ablation process to achieve a desiredgeometry. Because trench 142 provides control of the width of thefeature, no such trimming step is needed in the process 150 of FIG. 12.

FIG. 11 also illustrates a second trench 160 formed in the TBC material134 to a second depth that is farther below surface 138 than trench 142.By forming a plurality of trenches 142, 160 at a plurality of depthsbelow surface 138, it is possible to place sensors, such as sensors 50at more than one depth within the component 130, thereby furtheraugmenting the available operating parameter data. In the embodiment ofFIG. 11, trench 160 contains two vertically stacked conducting layers162, 164 separated by an insulating layer 166. The conducting layers162, 164 may form two portions of a sensor, or two conducting lines forconnecting a sensor to a connecting location. 1As illustrated in FIG.12, the two conducting layers 162, 164 may be formed by first depositingconducting layer 162 at step 158, and then depositing an insulatinglayer 166 at step 168 using any desired deposition technique, such asCVD.

Steps 158, 168 are then repeated to deposit conducting layer 164 andinsulating layer 174. The width of these layers is controlled by thewidth of trench 160 and the thickness of these layers may be controlledas they are deposited to achieve predetermined performancecharacteristics. For example, the thickness of insulating material 166will affect the impedance between the two conducting layers 162, 164.Conducting layer 164 is then isolated from the external environment bybackfilling the trench 160 with a barrier material such as thermallyinsulating material 170 at step 172. Insulating material 170 may be thesame material as TBC material 134 or a different material having desiredcharacteristics. Insulating material 170 may be deposited by any knowndeposition technique, including CVD, thermal spraying, selective lasermelting, or selective laser sintering. Selective laser melting andselective laser sintering processes are known in the art, as exemplifiedby Chapters 6 and 7 of “Laser-Induced Materials and Processes For RapidPrototyping” by L. Lu, J. Y. H. Fuh, and Y. S. Wong, published by KluwerAcademic Publishers.

Additional sensors 176, 178 may be disposed at preselected depths withincomponent 130 by forming respective trenches 180, 182 to appropriatedepths. Trenches 180, 182 may be backfilled with insulating material 170to the level of surface 138 at step 172. Planarization of surface 138may be performed at step 184, if necessary, such as when surface 138forms part of an airfoil. By forming a trench to a desired depth, asensor may be embedded to within the TBC material layer 134, to withinthe bond coat material layer 136, to within the substrate material 132,or to a depth of an interface between any two of these layers.

Thus, it is possible to develop actual operating parameter data across adepth of a component or across the depth of the thermal barrier coating.Such data may be useful for confirming design assumptions and forupdating computerized models, and it may also be useful as an indicatorof damage or degradation of a TBC coating. For example, a sensor 178embedded below the TBC material 134 may produce a signal indicating asignificant temperature rise in the event of cracking or spalling of thelayer of TBC material 134. Alternatively, the detection of apredetermined level of vanadium, sodium or sulfur deposits by anembedded sensor 176 may announce conditions that would give rise tospalling and failure of the TBC coating 134 if the component were toremain in service for an extended period. This process facilitates theplacement of sensors at any location on a fully assembled and coatedpart. Electrochemical sensors on the component surface can play animportant role in determining the nature and effect of corrosionproducts present in the surrounding environment.

MEMS sensors or devices typically include microelectronic packaging,integrating antenna structures for command signals intomicroelectromechanical structures for desired sensing or actuationfunctions. Silicon and high temperature electro-ceramics, such as GaN,SiC and AlN micromaching as well as others are advanced micromachingtechnologies that are commonly used to fabricate MEMS devices havingdimensions in the sub-millimeter range. This allows for fashioningmicroscopic mechanical parts out of silicon substrate or on a siliconsubstrate, making the structures 3-dimensional, which allows for anarray of applications. Electronic circuits functioning as transmittersand antennas may also be imprinted on the chips for wirelesstransmission. The inventors of the present invention have determinedthat deploying MEMS devices as integral parts of various components andlocations of combustion turbine 10 allows for improved monitoring ofcomponent and system operating parameters. This allows for improveddiagnostics, predictive maintenance and proof of design. Anotheradvantage is prognosis for design, which may use physics-basedapproaches towards understanding failures.

FIG. 13 illustrates a component 200 that may be formed by depositing afirst sensor 210 onto a surface of a substrate 212. Subsequently, afirst layer 214 of a barrier coating 216, such as a CMC abradablecoating system disclosed in U.S. Pat. No. 6,197,424, for example, isdeposited over the sensor 210. A second sensor 220 is then depositedover the first layer 214. A second layer 218 of barrier coating 216 isthen deposited, followed by the deposition of a third sensor 222 andthird layer 224 of the barrier coating. In this manner, one or moresensors 210, 220, 222, which may be various MEMS devices configured forperforming various functions may be embedded at a plurality of depthswithin the confines of a wall of a component 200. One may appreciatethat the same component 200 may be formed with various combinations ofMEMS sensors 210, 220, 222 configured to monitor various types ofconditions associated with component 200.

For example, embodiments of the structure of FIG. 13 may be useful formonitoring various properties of coating 216 such as the amount of wearof an abradable coating system, since each of the sensors 210, 220, 222may become exposed at a different time as the coating 216 undergoes weardue to abrasion. Signals generated by the respective sensors 210, 220,222 may be responsive to the wear of coating 216 and may be used in animproved clearance control program for predicting the remaining usefullife of an abradable coating and/or for estimating the amount of leakagepast an abradable seal.

FIG. 14 illustrates a schematic plan view of an exemplary MEMS device orsensor 250 that may be affixed to a component's substrate, such asdirectly onto a surface of the substrate, using phosphate cement orglue. MEMS device 250 may be affixed beneath the substrate's surfacesuch as by affixing within an indentation or recess then covered with anover layer of protective coating. It may alternately be affixed to thesubstrate by being retained within a barrier coating or otherwiseproperly secured in place for its intended purpose. Embodiments of MEMSsensor 250 may be conductively coupled to leads 258, 260, 266, 268 andconductors 262, 264, 270, 272, respectively, which may be depositedusing thermal spray deposition, for example, such as the conformaldirect write technology disclosed in U.S. Pat. No. 6,576,861. Otherdeposition processes may be used as recognized by those skilled in theart.

A plurality of sensors 250 may be affixed at varying depths within acoating such as an abradable coating system 216. Thermal sprayedabradable coating systems 216 are typically applied for gas pathclearance control, which influences power output and efficiency ofcombustion turbine 10. Coating systems 216 are usually porous coatingsthat abrade when contacted by a moving structural component, such as thetips of blades 18 and are designed not to damage the contacting surface.Information with respect to the wear behavior of coating system 216 maybe used to predict the useful life of the coating, prevent catastrophicinteraction between components and allow for improved control ofcombustion turbine 10.

MEMS sensor 250 may be configured as a proximity sensor that operatesunder capacitance or inductance. Sensor 250 may be an inductiveproximity sensor comprising a coil, an oscillator, a detection circuitand an output circuit as recognized by those skilled in the art. Theoscillator generates a fluctuating magnetic field around the winding ofthe coil that locates in the MEMS device's sensing face. When a metalobject moves into the inductive proximity sensor's field of detection,eddy circuits build up in the metallic object, magnetically push back,and finally dampen the sensor's 250 own oscillation field. The detectioncircuit monitors the oscillator's strength and triggers an output fromthe output circuitry when the oscillator becomes dampened to asufficient level.

One or more conductive connectors, such as a connector 52 may beprovided as a means for routing data signals indicative of the measuredresponse from sensor 250 to a transmitter 54, which may be configuredfor wirelessly transmitting the data signal to a transceiver 56, such asthose shown in FIG. 1. Connector 52 may be one or a plurality ofelectrical leads for conducting a signal from sensor 250 via conductors270, 272 to a transmitter such as surface mounted transmitter 54.Alternate embodiments allow for the signal to be conducted to an antenna(not shown), which may be an inductively couple spiral coil forwirelessly transmitting data signals from sensor 250 to a transmitter 54and/or transceiver 56.

Exemplary embodiments of MEMS sensor 250 may be configured to produce aneddy current circuit to detect intrusions into abradable coating system216. Such intrusions may be the tips of rotating blades 19 in compressor12 or blades 18 in turbine 16 abrading coating 216 during operation ofcombustion turbine 10. Intrusions between other components may bedetected within the casing of compressor 12 or turbine 16 at variousother places of interest.

Embodiments of MEMS sensor 250 may be an inductive proximity sensorhaving circuitry that generates an electromagnetic field and detects anychanges in a resonant circuit caused by eddy current losses induced in aconductive material influencing the magnetic field. When an AC voltageis applied to MEMS sensor 250 an oscillating current radiates anelectromagnetic field. When an electrical conductor or metal componentsuch as a tip of blade 18, for example, enters the electromagneticfield, eddy currents are drawn from the oscillator and induced into theblade tip. The losses in energy caused by the eddy currents may becorrelated to the distance and position of the blade tip relative toMEMS sensor 250.

FIG. 15 is a partial perspective view of turbine blades 18 intrudinginto abradable coating system 216 during rotation of the blades such aswhen combustion turbine 10 is in operation. A plurality of blades 18 ismounted to a rotor disk 280. Blade tips 282 are located just inside aninner wall 284, which may be a blade outer air seal or ring segment ofcombustion turbine 10 as recognized by those skilled in the art.Abradable coating system 216 may be deposited on a ring segment 284 sothat a groove 286 is abraded within the coating as blades 18 rotate. Oneor more MEMS sensors 250 may be affixed on or within the inner surfaceof ring segment 284, or within abradable coating system 216.

MEMS sensors 250 are depicted schematically as boxes in FIG. 15 but itwill be appreciated they may be configured to perform various functionsand be affixed in various configurations, orientations and locations.Embodiments of the invention may be used for continuously measuring thedistance between blade tip 282 and one or more MEMS sensors 250 duringoperation of combustion turbine 16. In this aspect, a first distancebetween the end of a blade tip 282, or other selected locations on ablade 18, and the location of one or more MEMS sensors 250 is known. Thefirst distance may be calculated and stored in database 36 of monitoringsystem 30 and may be the distance between a blade tip 282 and a MEMSsensor 250 prior to the commissioning of a combustion turbine 10. Thefirst distance may be other distances depending on the desiredmeasurements to be taken. It will be appreciated that blade tip 282 maybe coated with a barrier coating such as TBC 26 (FIG. 1). Thecomposition and thickness of such a coating may be accounted for whenselecting a location for one or more MEMS sensors 250 and calculatingwear of coating system 216.

Abradable coating system 216 has a first or original thickness wheninitially deposited on ring segment 284, or after repairing the coating,and prior to being abraded by blade tips 282. One or a plurality of MEMSsensors 250 may be affixed within coating system 216 at varying selecteddepths from the original surface 288 of coating system 216. Forinstance, a plurality of MEMS sensors 250 may be affixed in spacedrelation around the circumference of ring segment 284 for taking arespective plurality of measurements with respect to a single row ofblades 18. Ring segment 284 may include a row of ring segment sectionsthat circumscribe the row of blades. Using MEMS sensors 250 isadvantageous because a large quantity may be affixed in an area ofinterest to ensure data extraction in the event one or more sensorsfail.

As blade tips 282 of the row of blades 18 abrade coating system 216,groove 286 is formed within coating 216 that is approximately the widthof blades 18. Blades 18 abrading coating 216 forms a second or operatingthickness of that portion of coating system 216 that is not worn away byblade tips 282. This operating thickness may be defined as the thicknessof coating system 216 from the surface of groove 286 to the interface283 of coating system 216 with ring segment 284. The operating thicknessmay vary around the circumference of a respective ring segment 284 asappreciated by those skilled in the art.

The plurality of MEMS sensors 250 may continuously transmit data tomonitoring system 30 indicative of the distance between a respectivesensor 250 and a respective blade tip 282. Data indicative of thedistances blade tips 288 have traveled into coating 216 may be stored indatabase 36. This data may be used by processor 34 to calculate theamount or depth of wear the abradable coating system 216 is experiencingaround the circumference of ring segment 284. In this respect, processor34 may calculate the distance one or more blade tips 282 have traveledinto coating system 216 during operation of combustion turbine 10 suchas when going from start-up to full load.

Processor 34 may calculate the size of gaps formed between a blade tip282 and the inner surface of groove 286, including its edges, such asgaps formed when a blade tip 282 contracts from its maximum incursioninto coating system 216. Such gaps may be calculated knowing theoriginal thickness of coating 216, the maximum incursion of blade tips288 into coating 216 and the current distance between MEMS sensors 250and blade tips 288. This allows for estimating secondary gas path flowpast through the gaps, which may be used for more efficient operation ofcombustion turbine 10 and improved predictive maintenance. Calculationsmade by processor 34 based on data from MEMS sensors 250 may be relatedto operating cycles of combustion turbine 10 for various purposesincluding improved control during operating, cooling and service cyclesof combustion turbine 10, and avoidance of catastrophic failure.

Components within compressor 12 and turbine 16 may have different ratesof thermal expansion so they expand and contract at different ratesduring heating and cooling of turbine 16. Blades 18 may expand morequickly than a rotor to which rotor disk 280 is mounted due todifferences in their shape and mass. A control module of system 30 mayuse real-time and historical data from MEMS sensors 250 during operationor a heating and/or cooling cycle of turbine 16 to prevent blade tips282 from impinging on the inner surface of ring segment 284 bycontrolling various operating parameters of combustion turbine 10. Forexample, the turbine engine ramp rates and shut down schedule as well asscheduled spin cool cycles may be controlled in response to datareceived from MEMS sensors 250.

This data may also be used to control combustion turbine 10 to avoidother “pinch points”, which may occur between numerous components withincompressor 12 or turbine 16 during heating and/or cooling cycles. Such“pinch points” may develop for numerous reasons such as distortion ofring segment 284 due to servicing, uneven wear around ring segment 284,or the encroachment of ring segment 284 toward blade tips 282. This mayhappen due to vibration-induced wear on the hook portions of the ringsegment holding it in place within an isolation ring.

Aspects of the invention allow for using various embodiments of MEMSsensors 250 to directly interrogate components and coatings withincombustion turbine 10 to acquire data indicative of information that isa function of the type of MEMS sensor used. MEMS devices configured toperform more than one function may also be used. FIG. 14 is a schematicof a MEMS device 250 that may be configured to perform various functionssuch as an accelerometer, for example, that may be used to measurevibration. An exemplary MEMS accelerometer 250 may consist of a proofmass suspended by a spring such as a cantilever or beam. When MEMSaccelerometer 250 is subjected to acceleration, the inertia of the masscauses changes in the gap between it and the bulk of the device. Theprinciple of measuring the gap between the mass and bulk of the devicecan be performed using a capacitive, piezoresistive, piezoelectric,thermal, resonance or surface acoustic waves (SAW) principle asrecognized by those skilled in the art.

In this respect, response data from a MEMS accelerometer 250 may beanalyzed by monitoring system 30 to determine vibration frequency,forces and displacement of components within combustion turbine 10. Thisinformation may be used in assessing operating conditions of combustionturbine 10 including setting air and fuel maps during commissioning,monitoring for wear as a result of service, evaluating combustiondynamics, evaluating components for changes in natural frequency, whichmay be indicative of cracks or other defects, and validation of designmethodologies.

Embodiments of MEMS sensor 250 may be configured to perform otherfunctions such as a load cell, pressure transducer, strain gauge, ortemperature and heat flux sensors, for example, as recognized by thoseskilled in the art. A MEMS load cell 250 may include two bonded siliconwafers where the bottom layer contains an electrode pattern forming anarray of capacitors with the top wafer acting as a common electrode. Theload may be estimated through a change in capacitance between theflexible electrode and the rigid electrode. MEMS load cell 250 may beused to assess boundary conditions between components. Specifically,both static and dynamic contact force between two components may bemeasured. Such measurements may be used by monitoring system 30 toassess surface bearing stresses, critical in the prevention of wear, aswell as providing feedback for calibration and validation of designmethodologies and boundary conditions.

Embodiments allow for using MEMS pressure sensors 250, such as those inthe two general classes of: (a) piezoresistive where a silicon diaphragmconsisting of a few resistors in a Wheatstone bridge configurationallows for sense changes in pressure through changes in resistance; and(b) capacitive where the capacitance between a flexible membrane and afixed plate changes as a function of pressure. MEMS pressure transducers250 may be deployed in various places within combustion turbine 10 suchas the engine inlet for measurement of pressure distribution ordetection of inlet pressure instabilities. Also, using MEMS pressuresensors 250 for sensing between stages of compressor 12 allows fordetecting rotating stalls and early surge detection. This data may beused by monitoring system 30 for controlling operation of combustionturbine 10.

Embodiments allow for using MEMS devices 250 configured as strain gauge,temperature and heat flux sensors that may be based on the principlesdisclosed in “Microsensors, microelectromechanical systems (MEMS), andelectronics for smart structures and systems” by V. K. Varadan and V. V.Varadan published by IOP Publishing LTD, United Kingdom. Such MEMSdevices 250 may utilize the surface acoustic wave (SAW) properties ofmaterials to measure static and dynamic strain, and the thermoelectricproperties of the materials to measure temperature and heat flux. A SAWis typically a piezoelectric wafer such as lead zirconium titanate (PZT)and lithium niobate (LiNbO₅) class of materials, with interdigitaltransducers (IDT) and reflectors on its surface. The IDT convertselectrical energy into mechanical energy and vice versa for generatingand detecting SAW. The temperature and strain sensitive properties ofthe above class of materials may further be taken advantage of inmeasuring temperature and heat flux.

Embodiments of MEMS devices 250 configured as strain gauges, temperatureand heat flux sensors may be deployed in various places withincombustion turbine 10. For example, MEMS strain sensors 250 may beaffixed on blades 18, 19 in compressor 12 and turbine 16 for measuringstatic and dynamic strains. MEMS temperature and heat flux sensors 250may be utilized on blades 18 and vanes 22 in turbine 16 for measuringcomponent thermal state, such as hot spots, cooling effectiveness,thermal efficiency and heat flux transients within the component orcoating. This allows for an improved understanding of thermalenvironments in turbine 16 for materials development and designvalidation.

MEMS sensors 250 may be embedded directly into the surface of acomponent or frame of combustion turbine 10 as well as within coatingsdeposited on the component or frame. The MEMS sensors 250 may beinsulated from the surrounding component or frame via an insulatinglayer of material that may be deposited by thermal spray or othertechniques. Appropriate electrical connections may be made usingconventional wires or conductive leads deposited by microplasma spraytechnology or other techniques such as ones described herein. Overlayers of coatings may be deposited as necessary such as ones for wearresistance, dimensional control, oxidation resistance and thermalbarriers.

While the preferred embodiments of the present invention have been shownand described herein, it will be obvious that such embodiments areprovided by way of example only. Numerous variations, changes andsubstitutions will occur to those of skill in the art without departingfrom the invention herein. Accordingly, it is intended that theinvention be limited only by the spirit and scope of the appendedclaims.

1. A component for use in a combustion turbine, the componentcomprising: a substrate; a MEMS device affixed to the substrate; and atleast one connector in electrical communication with the MEMS device forrouting a data signal from the MEMS device to a termination location. 2.The component of claim 1 further comprising: a transmitter in electricalcommunication with the at least one connector for wirelesslytransmitting the data signal outside the combustion turbine.
 3. Thecomponent of claim 1 further comprising a barrier coating deposited onthe substrate wherein the MEMS device is affixed beneath a surface ofthe barrier coating.
 4. The component of claim 3 further comprising atrench formed in the barrier coating wherein the MEMS device is affixedto the barrier coating within the trench.
 5. The component of claim 4further comprising a backfill material deposited in the trench over theMEMS device.
 6. The component of claim 3 further comprising a pluralityof trenches formed in the barrier coating at respective different depthsbelow the surface of the barrier coating and a MEMS device depositedwithin each of the plurality of trenches.
 7. The component of claim 3,the MEMS device comprising a proximity sensor configured to detectintrusion of an object into the barrier coating.
 8. The component ofclaim 1, the MEMS device selected from the group consisting of a MEMSaccelerometer, a MEMS load cell, a MEMS strain gauge, a MEMS temperaturesensor, a MEMS heat flux sensor, a MEMS flow sensor and a MEMS pressuretransducer.
 9. A turbine comprising: a turbine comprising a row ofturbine blades and a row of ring segments circumscribing the row ofblades; a MEMS device affixed to a component within the turbine; atleast one connector in electrical communication with the MEMS device forrouting a data signal from the MEMS device to a termination location; amonitoring system comprising a processing module programmed forreceiving data from the MEMS device; and a transmitter for receiving thedata signal from the termination location and transmitting the datasignal to the monitoring system.
 10. The turbine of claim 9, the MEMSdevice selected from the group consisting of a MEMS accelerometer, aMEMS load cell, a MEMS strain gauge, a MEMS temperature sensor, a MEMSheat flux sensor, a MEMS flow sensor and a MEMS pressure transducer. 11.The turbine of claim 9 further comprising a barrier coating deposited ona ring segment wherein the MEMS device is affixed beneath a surface ofthe barrier coating.
 12. The turbine of claim 11 further comprising atrench formed in the barrier coating wherein the MEMS device is affixedto the barrier coating within the trench.
 13. The turbine of claim 9further comprising: an abradable coating system deposited on a ringsegment; the MEMS device comprising a proximity sensor affixed beneath asurface of the abradable coating system configured to detect intrusionof an object into the abradable coating system; and wherein theprocessing module is programmed to calculate a clearance between a tipof at least one blade of the row of blades and the MEMS device based ondata received from the MEMS device.
 14. The turbine of claim 9 furthercomprising: a barrier coating deposited on the component; the MEMSdevice comprising a pressure transducer affixed beneath a surface of thebarrier coating; and wherein the processing module is programmed todetect wear of the barrier coating based on data received from the MEMSdevice.
 15. The turbine of claim 9 further comprising: a barrier coatingdeposited on the component; and a plurality of trenches formed in thebarrier coating at respective different depths below the surface of thebarrier coating and a MEMS device deposited within each of the pluralityof trenches.
 16. The turbine of claim 9, the MEMS device comprising aload cell configured to measure static and dynamic contact forcesbetween the first component and a second component within the turbine.17. The turbine of claim 9, the MEMS device comprising a strain gaugeaffixed to at least one of the plurality of turbine blades configured tomeasure static and dynamic strains on the at least one of the pluralityof turbine blades.
 18. The turbine of claim 9, the MEMS devicecomprising a heat flux sensor affixed to the first component configuredto measure a thermal state of the component.
 19. The turbine of claim 9further comprising: an abradable coating system deposited on a surfaceof the ring segment; a trench formed in the abradable coating system;the MEMS device deposited in the trench; and a backfill materialdeposited in the trench over the MEMS device.
 20. A monitoring systemfor use with a turbine, the monitoring system comprising: at least oneMEMS device configured for use with the turbine; a processing moduleprogrammed to receive data from the at least one MEMS device and todevelop information based on the received data; and a transmitter fortransmitting data from the at least one MEMS device to the processingmodule.
 21. The monitoring system of claim 20, the processing moduleprogrammed to generate a control signal responsive to the received data,the control signal useful for controlling an operating parameter of theturbine.
 22. The monitoring system of claim 21, the at least one MEMSdevice comprising a MEMS proximity sensor wherein the control signal isuseful for controlling a clearance between a row of turbine blades and aring segment.
 23. The monitoring system of claim 20, the at least oneMEMS device comprising a load cell wherein the processing module isprogrammed to develop information with respect to static and dynamiccontact forces between a first component within the turbine and a secondcomponent within the turbine.
 24. The monitoring system of claim 20, theat least one MEMS device comprising a heat flux sensor wherein theprocessing module is programmed to develop information with respect to athermal state of a component within the turbine.
 25. The monitoringsystem of claim 20, the at least one MEMS device comprising a straingauge wherein the processing module is programmed to develop informationwith respect to static and dynamic strains on at least one turbineblade.